Non-contacting seals for geared gas turbine  engine bearing compartments

ABSTRACT

A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section. The engine also includes a rotating element and at least one bearing compartment including a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. A method and section for a gas turbine engine are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.15/402,316, filed Jan. 10, 2017, which is a continuation in part of U.S.application Ser. No. 14/243,003, filed Apr. 2, 2014, which is acontinuation of U.S. application Ser. No. 14/053,648, filed Oct. 15,2013, which is a continuation of U.S. application Ser. No. 13/787,919,filed Mar. 7, 2013.

BACKGROUND

Gas turbine engines are known and, when utilized in aircraftapplications, typically include a fan delivering air into a bypass ductand into a core engine flow. The core engine flow passes into acompressor where the air is compressed and then delivered into acombustion section. The air is mixed with fuel in the combustion sectionand ignited. Products of that combustion pass downstream over turbinerotors, driving them to rotate.

Historically, a fan drive turbine drove the fan through a direct drive,such that they rotated at the same speed. This restricted the speedavailable for the fan drive turbine, as the fan speed was limited.

More recently, it has been proposed to include a gear reduction betweenthe fan drive turbine and the fan. With this change, the speed of thefan drive turbine can increase.

In gas turbine engines, there are a number of bearing compartments whichare desirably sealed. In the prior art, operating at slower speeds,contact seals have been utilized, which directly contacted surfacesrotating with the shaft to seal the bearing compartments. Such contactseals were typically cooled using oil or other lubricant, which wascirculated through a cooling system. For geared engines, in whichcertain components are enabled to rotate faster than correspondingcomponents in non-geared engines, to achieve the same amount of coolinga larger volume of lubricant would be needed. Moreover, a larger volumeof lubricant would require a larger holding tank and correspondinglylarger cooling system fluid pumping apparatus. All of the larger volumeof lubricant, the larger holding tank, and the larger fluid pumpingapparatus would add undesirable weight to the engine.

SUMMARY

A seal arrangement for a bearing compartment of a gas turbine engineaccording to an example of the present disclosure includes a rotatingelement and at least one bearing compartment that has a bearing forsupporting the rotating element. The rotating element is rotatable aboutan axis. The bearing compartment has a first seal and a second seal eachassociated with a corresponding one of two opposed axial ends, on eitheraxial side of the bearing relative to the axis. The first seal and thesecond seal are a non-contacting seal that have a seal face facing arotating face of the rotating element. A radially outermost location ofthe bearing compartment defines a compartment radius with respect to theaxis. A radial outermost location of the non-contacting seal establishesa seal radius with respect to the engine axis, and a compartment-sealratio defined by the compartment radius to the seal radius is less thanor equal to 6:1.

In a further embodiment of any of the foregoing embodiments, thenon-contacting seal is arranged to resist leakage of lubricant outwardlyof the at least one bearing compartment and to allow pressurized air toflow from a chamber adjacent the non-contacting seal into the at leastone bearing compartment, and a grooved area is formed in one of thefaces, with the grooved area having a plurality of circumferentiallyspaced grooves for generating hydrodynamic lift-off forces and allowingleakage of pressurized air across the faces and into the at least onebearing compartment to resist leakage of lubricant from the at least onebearing compartment.

In a further embodiment of any of the foregoing embodiments, therotating element is configured to rotate at a velocity greater than orequal to 450 feet per second.

In a further embodiment of any of the foregoing embodiments, thenon-contacting seal is formed with a plurality of passages configured toallow tapping of additional pressurized air to be delivered to the facesat a location in the proximity of the grooved area for generatinghydrostatic lift-off forces.

In a further embodiment of any of the foregoing embodiments, the groovedarea is spaced radially from the plurality of passages at the seal face.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotatable with a rotor having an axial facefacing the seal face, and the grooved area is formed in the rotor.

In a further embodiment of any of the foregoing embodiments, thenon-contacting seal is a controlled gap carbon seal having a full hoopseal and a metal band shrunk fit onto the non-contacting seal, andpositioned in a seal carrier.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotatable with a rotor having an axial facefacing the seal face.

In a further embodiment of any of the foregoing embodiments, therotating element is configured to rotate at a velocity greater than orequal to 450 feet per second.

In a further embodiment of any of the foregoing embodiments, thecompartment-seal ratio is between 3:1 to 5:1.

In a further embodiment of any of the foregoing embodiments, therotating element is configured to rotate at a velocity less than orequal to 600 feet per second.

A gas turbine engine according to an example of the present disclosureincludes a fan section having a plurality of fan blades, a geararrangement, a compressor section, and a turbine section arranged alongan engine axis. The turbine section includes a first turbine and asecond turbine. The second turbine is configured to drive the fansection through the gear arrangement. A seal arrangement includes arotating element and at least one bearing compartment that has a bearingfor supporting the rotating element. The bearing compartment has a firstseal and a second seal each associated with a corresponding one of twoopposed axial ends, on either axial side of the bearing relative to theengine axis, at least one of the first seal and the second seal is anon-contacting seal that has a seal face facing a rotating face of therotating element. A radially outermost location of the fan bladesdefines a fan radius with respect to the engine axis. A radiallyoutermost location of the bearing compartment defines a compartmentradius with respect to the engine axis, and a fan-compartment ratiodefined by the fan radius to the compartment radius is greater than orequal to 2:1.

In a further embodiment of any of the foregoing embodiments, the firstturbine is configured to drive the rotating element.

In a further embodiment of any of the foregoing embodiments, a radiallyoutermost location of the bearing compartment defines a compartmentradius with respect to the engine axis, a radial outermost location ofat least one of the first and second seals establishes a seal radiuswith respect to the engine axis, and a compartment-seal ratio defined bythe compartment radius to the seal radius is less than or equal to 6:1.

In a further embodiment of any of the foregoing embodiments, therotating element is configured to rotate at a velocity greater than orequal to 450 feet per second.

In a further embodiment of any of the foregoing embodiments, a groovedarea is formed in one of the faces, with the grooved area having aplurality of circumferentially spaced grooves for generatinghydrodynamic lift-off forces and allowing leakage of pressurized airacross the faces and into the at least one bearing compartment to resistleakage of lubricant from the at least one bearing compartment.

In a further embodiment of any of the foregoing embodiments, therotating element is configured to rotate at a velocity between 450 feetper second and 600 feet per second.

A method of operating a gas turbine engine according to an example ofthe present disclosure includes the steps of arranging a bearing withina bearing compartment to support a rotating element. The rotatingelement defines a rotating face. The bearing compartment has a firstseal and a second seal each associated with a corresponding one of twoopposed axial ends relative to an engine axis, on either axial side ofthe bearing. The method includes rotating the rotating face relative toat least one of the first seal and the second seal, and sealing thebearing compartment with the first seal and the second seal. The firstseal and the second seal are a non-contacting seal configured to resistleakage of lubricant outwardly of the bearing compartment and to allowair to flow from a chamber adjacent the non-contacting seal and into thebearing compartment. The non-contacting seal defines a seal face facingthe rotating face. A radially outermost location of the bearingcompartment defines a compartment radius with respect to the engineaxis. A radial outermost location of the non-contacting seal establishesa seal radius with respect to the engine axis, and a ratio of thecompartment radius to the seal radius is less than or equal to 6:1.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotatable with a rotor having an axial facefacing the seal face.

A further embodiment of any of the foregoing embodiments includescommunicating air from a fan to a bypass passage and to a compressorsection. A bypass ratio is defined as the volume of air passing into thebypass passage compared to the volume of air passing into the compressorsection. The bypass ratio is greater than 10 at a cruise condition. Thestep of rotating includes rotating the rotating element at a velocitygreater than or equal to 450 feet per second, with the rotating elementdriving the compressor section.

A gas turbine engine according to an example of the present disclosureincludes a fan, a compressor section, a combustor, and a turbinesection, a rotating element and at least one bearing compartmentincluding a bearing for supporting the rotating element, a seal forresisting leakage of lubricant outwardly of the bearing compartment andfor allowing pressurized air to flow from a chamber adjacent the sealinto the bearing compartment. The seal has a seal face facing a rotatingface rotating with the rotating element, and the seal is a non-contactseal. The bearing compartment has a seal associated with each of twoopposed axial ends on either axial side of the bearing.

In a further embodiment of the foregoing embodiment, a grooved area isformed in one of the faces. The grooved area has a plurality ofcircumferentially spaced grooves for generating hydrodynamic lift-offforces and allows leakage of pressurized air across the faces and intothe bearing compartment to resist leakage of lubricant from the bearingcompartment.

In a further embodiment of either of the foregoing embodiments, the sealis formed with a plurality of passages to allow tapping of additionalpressurized air to be delivered to the faces at a location in theproximity of the grooved area for generating hydrostatic lift-offforces.

In a further embodiment of any of the foregoing embodiments, the groovedarea is spaced radially from the plurality of passages at the seal face.

In a further embodiment of any of the foregoing embodiments, each of theplurality of passages is positioned radially outward of the groovedarea.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotating with a rotor having an axial facefacing the seal face.

In a further embodiment of any of the foregoing embodiments, the groovedarea is formed in the rotor.

In a further embodiment of any of the foregoing embodiments, the turbinesection includes a fan drive turbine driving the fan through a gearreduction. The rotating element is driven by the fan drive turbine. Atleast one bearing compartment is associated with the gear reduction.

In a further embodiment of any of the foregoing embodiments, the seal isa carbon seal.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotating with a rotor having acircumferential face facing the seal face.

In a further embodiment of any of the foregoing embodiments, the sealface faces radially inwardly.

In a further embodiment of any of the foregoing embodiments, a groovedarea is formed in one of the faces, with the grooved area having aplurality of circumferentially spaced grooves for generatinghydrodynamic lift-off forces and allowing leakage of pressurized airacross the faces and into the bearing compartment to resist leakage oflubricant from the bearing compartment.

In a further embodiment of any of the foregoing embodiments, the groovedarea is formed in the rotor.

In a further embodiment of any of the foregoing embodiments, the seal isa circumferentially segmented carbon seal.

In a further embodiment of any of the foregoing embodiments, the seal isa controlled gap carbon seal having a full hoop seal and a metal bandshrunk fit onto the seal, and positioned in a seal carrier.

In a further embodiment of any of the foregoing embodiments, therotating element is driven by a fan drive turbine. At least one bearingcompartment is associated with a gear reduction for driving the fan.

A method of designing a section of a gas turbine engine according to anexample of the present disclosure includes configuring a bearingcompartment to include a bearing designed to support a rotating element,configuring the rotating element to define a rotating face, the rotatingface configured to rotate with said rotating element, configuring thebearing compartment to include a seal designed to resist leakage oflubricant outwardly of the bearing compartment and to allow air to flowfrom a chamber adjacent the seal and into the bearing compartmentconfiguring the seal to define a seal face facing the rotating face,designing the seal to be a non-contact seal, and configuring the bearingcompartment to have a seal associated with each of two opposed axialends, on either axial side of said bearing.

A further embodiment of the foregoing embodiment includes the step ofdesigning the faces to define a grooved area in one of the faces. Thegrooved area has a plurality of circumferentially spaced grooves forgenerating hydrodynamic lift-off forces and allows leakage ofpressurized air across the faces and into the bearing compartment toresist leakage of lubricant from the bearing compartment.

In a further embodiment of either of the foregoing embodiments, therotating element is designed to be a shaft rotating with a rotor havingan axial face facing the seal face.

In a further embodiment of any of the foregoing embodiments, the groovedarea is formed in the rotor.

A further embodiment of any of the foregoing embodiments includes thestep of designing the seal to define a plurality of passages to allowtapping of additional pressurized air to be delivered to the faces at alocation in the proximity of the grooved area for generating hydrostaticlift-off forces.

In a further embodiment of any of the foregoing embodiments, therotating element is designed to be a shaft rotating with a rotor havinga circumferential face facing the seal face.

In a further embodiment of any of the foregoing embodiments, the seal isdesigned to be a controlled gap carbon seal having a full hoop seal anda metal band shrunk fit onto the seal, and positioned in a seal carrier.

A section for a gas turbine engine according to an example of thepresent disclosure includes a rotating element and at least one bearingcompartment configured to be secured to a static structure. The bearingcompartment includes a bearing for supporting the rotating element and aseal for resisting leakage of lubricant outwardly of the bearingcompartment and for allowing pressurized air to flow from a chamberacross the seal into the bearing compartment. The seal has a seal facefacing a rotating face rotating with the rotating element. The seal is anon-contact seal where the bearing compartment has a seal associatedwith each of two opposed axial ends, on either axial side of thebearing.

In a further embodiment of the foregoing embodiment, a grooved area isformed in one of the faces. The grooved area has a plurality ofcircumferentially spaced grooves for generating hydrodynamic lift-offforces and allows leakage of pressurized air across the faces and intothe bearing compartment to resist leakage of lubricant from the bearingcompartment.

In a further embodiment of either of the foregoing embodiments, the sealis formed with a plurality of passages to allow tapping of additionalpressurized air to be delivered to the faces at a location in theproximity of the grooved area for generating hydrostatic lift-offforces.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotating with a rotor having an axial facefacing the seal face.

In a further embodiment of any of the foregoing embodiments, therotating element is a shaft rotating with a rotor having acircumferential face facing the seal face.

In a further embodiment of any of the foregoing embodiments, the seal isa circumferentially segmented carbon seal.

In a further embodiment of any of the foregoing embodiments, the seal isa controlled gap carbon seal having a full hoop seal and a metal bandshrunk fit onto the seal, and positioned in a seal carrier.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2A schematically shows example locations of bearing compartments.

FIG. 2B schematically shows dimensional aspects of the bearingcompartments of FIG. 2A.

FIG. 3A is a first embodiment of a non-contact seal according to thepresent invention.

FIG. 3B shows a second embodiment of a non-contact seal according to thepresent invention.

FIG. 3C shows a third embodiment of a non-contact seal according to thepresent invention.

FIG. 3D shows a fourth embodiment of a non-contact seal assemblyaccording to the present invention.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. A radially outermost location of thefan blades of the fan 42 establishes a radius R_(F) relative to theengine central longitudinal axis A. In some embodiments, radius R_(F) isbetween about 28 inches and about 40.5 inches. In an embodiment, theradius R_(F) is about 36.5 inches. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. A mid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and geared architecture 48 may be varied. For example,geared architecture 48 may be located aft of combustor section 26 oreven aft of turbine section 28, and fan section 22 may be positionedforward or aft of the location of geared architecture 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2A shows an embodiment of an arrangement of bearing compartments100 associated with the gas turbine engine, such as the gas turbineengine 20 illustrated in FIG. 1. As shown, a bearing compartment 102 isassociated with a low speed shaft 92 at a location associated with thelow pressure turbine. Bearings 106 are shown schematically as is a seal104.

A bearing compartment 108 is associated with a high speed rotor 90 andat the high pressure turbine of FIG. 1. Bearing compartment 108 includesseals 110 at each axial end and a central bearing 112.

Another bearing compartment 114 is also associated with the high speedrotor 90 and the high pressure compressor and includes a bearing 118 andseals 116.

Finally, a bearing compartment is associated with a fan drive gearsystem 122 at location 120 and with and the fan at location 123. Seals126 and 128 mechanically seal the axial ends of the bearing compartment120 and are associated with the fan rotor 127 and the low speed rotor92. The seals 126, 128 are also respectively associated with thebearings 124 and 130 that are positioned within the bearing compartment120/123.

Referring to FIG. 2B, the bearing compartments 102, 108, 114, 120/123,establish a dimensional relationship with respect to the engine centrallongitudinal axis A. A radially outermost location of bearingcompartments 102, 108, 114, 120/123 establish corresponding radii R₁₀₂,R₁₀₈, R₁₁₄ and R_(120/123).

Utilizing the seal arrangements disclosed herein, the relative sizes ofthe bearing compartments 102, 108, 114 and 120/123 can be reduced toprovide for a more compact core engine arrangement, which may beutilized with a relatively high bypass ratio. In some embodiments, radiiR_(120, 123) and R₁₁₄ establish a compartment ratio that is betweenabout 1:1 to about 2:1. In embodiments, radii R_(120, 123) and R₁₀₂establish a compartment ratio that is between about 2:1 to about 3:1. Insome embodiments, radii R₁₀₈ and R₁₀₂ establish a compartment ratio thatis between about 1:1 to about 2:1. For the purposes of this disclosure,the term “about” is relative to the number of significant digits unlessotherwise noted.

The sizes of the bearing compartments 102, 108, 114 and 120/123 relativeto the fan 42 can also be reduced. In embodiments, radii R_(F) and radiiR_(120/123) establish a fan-compartment ratio that is greater than orequal to about 2:1, such as between about 2.5:1 and about 5:1. In oneembodiment, the fan-compartment ratio is about 3:1. In embodiments,radii R_(F) and radii R₁₁₄ establish a fan-compartment ratio thatgreater than or equal to about 4:1. In an embodiment, radii R_(F) andradii R₁₁₄ establish a fan-compartment ratio that is about 5:1. Inembodiments, radii R_(F) and radii R₁₀₈ establish a fan-compartmentratio that greater than or equal to about 8:1, and less than or equal toabout 12:1. In an embodiment, radii R_(F) and radii R₁₀₈ establish afan-compartment ratio that is about 10:1. In embodiments, radii R_(F)and radii R₁₀₂ establish a fan-compartment ratio that is greater than orequal to about 12:1, and less than or equal to about 18:1. In anembodiment, radii R_(F) and radii R₁₀₂ establish a fan-compartment ratiothat is about 15:1.

There are challenges with sealing the bearing compartments in a gearedturbofan engine. Accordingly, various embodiments disclosed hereinrelate to the use of non-contacting seals such as lift-off seals at anyone or more of the locations of the seals shown in FIG. 2A or in anyother bearing compartment on a gas turbine engine. In some embodiments,the seals may be lift-off seals and, more particularly, may be carbonlift-off seals. Of course, in other embodiments other non-contactingseals, including other lift-off seals may be used.

Thus, as shown in FIG. 3A, a shaft 140, which could be any rotatingshaft in a gas turbine engine, has a mating rotor 142. An axial face 147of this mating rotor 142 is sealed relative to a face 145 from a seal144. The faces 145 and 147 face each other to form a mechanical seal.The seal 144 may be a non-contact seal such as a carbon seal lift-offseal. The interface between faces 145 and 147 experiences highvelocities, especially when compared to the prior art. The high velocityis a combination of a high rotational speed of the shaft 140 and arelatively large diameter for the seal 144. Velocities greater than orequal to about 450 feet/second (137.16 meters/second) may be seen.

In the FIG. 3A embodiment seal 144, a set of shallow grooves 152 isprovided by cutting into the face 147 of the rotor 142, as shown atcircumferentially spaced grooves 154. A spring 146 biases the seal 144toward the face 147. A higher pressure air is available in a chamber148, which is on an opposed side of the seal 144 from the bearingcompartment 150. The bearing compartment 150 is at a lower pressure thanthe chamber 148, and this higher pressure air passes through the groovedarea 152, such that the air flow levitates (lifts-off) the sealingsurface 145 of the non-rotating seal 144 from the sealing surface 147 ofthe rotor 142. The levitation is a result of hydrodynamic lifting forceas the air passes into the bearing compartment 150, preventing oil fromescaping the bearing compartment 150.

Another embodiment is illustrated in FIG. 3B. FIG. 3B provides themechanical sealing between face 168 of a non-rotating seal 162 and aface 170 of a rotor in a manner somewhat similar to the FIG. 3Aembodiment. There is a grooved area 172 having circumferentially spacedgrooves 174 with the features as described for the first embodiment thatgenerate hydrodynamic lifting force as the gas passes from a highpressure chamber 160 into the bearing compartment 274. Furthermore, thenon-rotating seal 162 has an inlet 166, a passage 164, and an outlet 180which delivers additional high pressure air generating hydrostaticlifting forces at a radial location in the proximity of the grooved area172, thereby providing a stronger and more stable lift-off seal comparedto the first embodiment. As shown, there is a plurality ofcircumferentially spaced outlets 180. The non-rotating seal 162, whichis biased toward the rotor 170 by a spring 161, may be a carbon lift-offseal.

FIG. 3C shows another embodiment 182, wherein the seal 186 has aplurality of circumferentially segmented members biased by spring 184toward a face 185 of a rotor 142 rotating with the shaft 140. Seal 186has a radially inwardly facing face 183 providing the seal face with themating face 185. One of the sealing faces, either 183 or 185, has a setof shallow, circumferentially spaced grooves 210 in a grooved area 211,somewhat similar to those described in the earlier embodiments thatgenerate a hydrodynamic force that levitates (lifts-off) thenon-rotating sealing surface 183 from the rotating mating surface 185when the high pressure chamber 188 delivers pressurized air across theseal 186 to prevent leakage of oil from the bearing compartment 190.

FIG. 3D shows an embodiment of a controlled gap non-contacting sealassembly 201. The shaft 140 has an outer surface spaced by a small gap196 from two carbon seals 192. The gap is controlled by design,typically by sizing the sealing diameters of the seals 192 and the rotor142 such that a small gap is maintained under all conditions. The shaftouter surface 193 and a radially inward facing surfaces 191 of the seals192 provide the seal faces. In one embodiment, the carbon seals 192 arefull hoop members extending around the entire circumference of the shaft140. A metal band 194 is shrunk fit onto the seal 192. A carrier 195mounts the seals 192. A high pressure chamber 198 is spaced from thebearing compartment 200, such that high pressure air passes through thegap 196 to prevent the leakage of lubricant.

The seal arrangements disclosed herein can be utilized to reduce therelative sizes of the bearing compartments to provide for a more compactcore engine arrangement. The seals 144, 162, 186, 192 establish adimensional relationship with engine central longitudinal axis A. Aradially outermost location of seals 144, 162, 186, 192 establishcorresponding radii R₁₄₄, R₁₆₂, R₁₈₆ and R₁₉₂ with respect to engineaxis A (shown in FIGS. 3A-3D, respectively). For the purposes of thisdisclosure, the radially outermost location of each seal 144, 162, 186,192 is measured at an outermost position that the seal 144, 162, 186,192 is permitted to move relative to the engine central longitudinalaxis A during normal operation of the engine 20. A radially outermostlocation of the bearing compartments 150, 274, 190, 200 establishcorresponding radii R₁₅₀, R₂₇₄, R₁₉₀ and R₂₀₀ with respect to engineaxis A (shown in FIGS. 3A-3D, respectively). For the purposes of thisdisclosure, the radially outermost location of each bearing compartment150, 274, 190, 200 is measured at an inner surface of the cavity of thebearing compartment 150, 274, 190, 200 housing the respectivebearing(s).

In embodiments, radii R_(150/274/190/200) and R_(144/162/186/192)establish a compartment-seal ratio that is less than or equal to about6:1. In some embodiments, the compartment-seal ratio is between about3:1 and about 5:1.

As previously discussed, the interface between the shaft or rotor andseal(s) may operate at relatively high velocities. In embodiments, theinterface between the faces of the shaft 140 or rotor 142 and seal(s)144/162/186/192 is configured to operate at relative velocities greaterthan or equal to about 450 feet/second. In some embodiments, theinterface operates at relative velocities less than or equal to about600 feet/second, and more narrowly less than or equal to about 550feet/second. In embodiments, the shaft 140 corresponds to the high speedrotor 90 (FIG. 2A).

In other embodiments, the interface between the faces of the shaft 140or rotor 142 and seal(s) 144/162/186/192 is configured to operate atrelative velocities less than or equal to about 450 feet/second, such asbetween about 400 to about 450 feet/second. In embodiments, the shaft140 corresponds to the low speed rotor 92 (FIG. 2A). In otherembodiments, the shaft 140 corresponds to the fan rotor 127 (FIG. 2A).In some embodiments, the interface between the faces of the shaft 140 orrotor 142 and seal(s) 144/162/186/192 is configured to operate atrelative velocities between about 200 to about 400 feet/second.

In some embodiments, a velocity of the high speed rotor 90 and the lowspeed rotor 92 establish a ratio that is between about 1:1 to about 2:1.In embodiments, a velocity of the high speed rotor 90 and the fan rotor127 establish a ratio that is between about 5:2 to about 3:1, with thefan rotor 127 driven by the fan drive gear system 122.

All of the disclosed embodiments reduce the friction between the sealand the rotating components. This reduces heat generation due tofriction, increases the durability of the seals, minimizes loss of oil,and increases the efficiency in fuel consumption of the overall engine.Moreover, as a result of the reduction in friction, less lubricant canbe used, thereby also reducing the size of the applicable fluid storagetank (not shown) and the applicable cooling system fluid pumpingapparatus (also not shown). Accordingly, the overall weight of theengine may be greatly reduced, thereby increasing the engine's fuelefficiency.

The disclosed embodiments may be useful at any bearing compartment in agas turbine engine. Although shafts are shown supported by the bearings,the disclosure would extend to other rotating elements supported by abearing.

Although various embodiments of this invention have been disclosed, aworker of ordinary skill in this art would recognize that certainmodifications would come within the scope of this invention. For thatreason, the following claims should be studied to determine the truescope and content of this invention.

1. A seal arrangement for a bearing compartment of a gas turbine enginecomprising: a rotating element and at least one bearing compartmentincluding a bearing for supporting said rotating element, said rotatingelement rotatable about an axis; wherein said at least one bearingcompartment has a first seal and a second seal each associated with acorresponding one of two opposed axial ends, on either axial side ofsaid bearing relative to said axis, at least one of said first seal andsaid second seal being a non-contacting seal having a seal face facing arotating face of said rotating element; and wherein a radially outermostlocation of said bearing compartment defines a compartment radius withrespect to said axis, a radial outermost location of said non-contactingseal establishes a seal radius with respect to the engine axis, and acompartment-seal ratio defined by said compartment radius to said sealradius is less than or equal to 6:1.
 2. The seal arrangement as setforth in claim 1, wherein said non-contacting seal is arranged to resistleakage of lubricant outwardly of said at least one bearing compartmentand to allow pressurized air to flow from a chamber adjacent saidnon-contacting seal into said at least one bearing compartment, and agrooved area is formed in one of said faces, with said grooved areahaving a plurality of circumferentially spaced grooves for generatinghydrodynamic lift-off forces and allowing leakage of pressurized airacross said faces and into said at least one bearing compartment toresist leakage of lubricant from said at least one bearing compartment.3. The seal arrangement as set forth in claim 2, wherein said rotatingelement is configured to rotate at a velocity greater than or equal to450 feet per second.
 4. The seal arrangement as set forth in claim 2,wherein said non-contacting seal being formed with a plurality ofpassages configured to allow tapping of additional pressurized air to bedelivered to said faces at a location in the proximity of said groovedarea for generating hydrostatic lift-off forces.
 5. The seal arrangementas set forth in claim 4, wherein said grooved area is spaced radiallyfrom said plurality of passages at said seal face.
 6. The sealarrangement as set forth in claim 5, wherein said rotating element is ashaft rotatable with a rotor having an axial face facing said seal face,and said grooved area is formed in said rotor.
 7. The seal arrangementas set forth in claim 2, wherein said non-contacting seal is acontrolled gap carbon seal having a full hoop seal and a metal bandshrunk fit onto said non-contacting seal, and positioned in a sealcarrier.
 8. The seal arrangement as set forth in claim 2, wherein saidrotating element is a shaft rotatable with a rotor having an axial facefacing said seal face.
 9. The seal arrangement as set forth in claim 1,wherein said rotating element is configured to rotate at a velocitygreater than or equal to 450 feet per second.
 10. The seal arrangementas set forth in claim 9, wherein said compartment-seal ratio is between3:1 to 5:1.
 11. The seal arrangement as set forth in claim 9, whereinsaid rotating element is configured to rotate at a velocity less than orequal to 600 feet per second.
 12. A gas turbine engine comprising: a fansection having a plurality of fan blades, a gear arrangement, acompressor section, and a turbine section arranged along an engine axis,said turbine section including a first turbine and a second turbine,said second turbine configured to drive said fan section through saidgear arrangement; a seal arrangement comprising: a rotating element andat least one bearing compartment including a bearing for supporting saidrotating element; wherein said at least one bearing compartment has afirst seal and a second seal each associated with a corresponding one oftwo opposed axial ends, on either axial side of said bearing relative tosaid engine axis, at least one of said first seal and said second sealbeing a non-contacting seal having a seal face facing a rotating face ofsaid rotating element; and wherein a radially outermost location of saidfan blades define a fan radius with respect to said engine axis, aradially outermost location of said bearing compartment defines acompartment radius with respect to said engine axis, and afan-compartment ratio defined by said fan radius to said compartmentradius is greater than or equal to 2:1.
 13. The gas turbine engine asset forth in claim 12, wherein said first turbine is configured to drivesaid rotating element.
 14. The gas turbine engine as set forth in claim12, wherein a radially outermost location of said bearing compartmentdefines a compartment radius with respect to said engine axis, a radialoutermost location of at least one of said first and second sealsestablishes a seal radius with respect to the engine axis, and acompartment-seal ratio defined by said compartment radius to said sealradius is less than or equal to 6:1.
 15. The gas turbine engine as setforth in claim 14, wherein said rotating element is configured to rotateat a velocity greater than or equal to 450 feet per second.
 16. The gasturbine engine as set forth in claim 15, wherein a grooved area isformed in one of said faces, with said grooved area having a pluralityof circumferentially spaced grooves for generating hydrodynamic lift-offforces and allowing leakage of pressurized air across said faces andinto said at least one bearing compartment to resist leakage oflubricant from said at least one bearing compartment.
 17. The gasturbine engine as set forth in claim 12, wherein said rotating elementis configured to rotate at a velocity between 450 feet per second and600 feet per second.
 18. A method of operating a gas turbine engine, themethod comprising the steps of: arranging a bearing within a bearingcompartment to support a rotating element, said rotating elementdefining a rotating face, said bearing compartment having a first sealand a second seal each associated with a corresponding one of twoopposed axial ends relative to an engine axis, on either axial side ofsaid bearing; rotating said rotating face relative to at least one ofsaid first seal and said second seal; sealing said bearing compartmentwith said first seal and said second seal, at least one of said firstseal and said second seal being a non-contacting seal configured toresist leakage of lubricant outwardly of said bearing compartment and toallow air to flow from a chamber adjacent said non-contacting seal andinto said bearing compartment, said non-contacting seal defining a sealface facing said rotating face; and wherein a radially outermostlocation of said bearing compartment defines a compartment radius withrespect to said engine axis, a radial outermost location of saidnon-contacting seal establishes a seal radius with respect to the engineaxis, and a ratio of said compartment radius to said seal radius is lessthan or equal to 6:1.
 19. The method as set forth in claim 18, whereinsaid rotating element is a shaft rotatable with a rotor having an axialface facing said seal face.
 20. The method as set forth in claim 18,comprising: communicating air from a fan to a bypass passage and to acompressor section, wherein a bypass ratio is defined as the volume ofair passing into said bypass passage compared to the volume of airpassing into said compressor section, said bypass ratio greater than 10at a cruise condition; and wherein said step of rotating comprisesrotating said rotating element at a velocity greater than or equal to450 feet per second, with said rotating element driving said compressorsection.